NREL's S805A Airfoil (s805a-nr) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: NREL's S805A Airfoil (s805a-nr) Reynolds number: 200,000 Max Cl/Cd: 66.98 at α=6.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-s805a-nr-200000-n5.txt Download as CSV file: xf-s805a-nr-200000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NREL's S805A Airfoil
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.200 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-11.000 -0.4529 0.09039 0.08680 -0.0487 1.0000 0.0160
-10.750 -0.4582 0.08545 0.08189 -0.0510 1.0000 0.0158
-10.500 -0.4712 0.07776 0.07425 -0.0559 1.0000 0.0157
-10.250 -0.4917 0.06989 0.06634 -0.0612 1.0000 0.0155
-10.000 -0.5145 0.06455 0.06095 -0.0630 1.0000 0.0153
-9.750 -0.5369 0.06069 0.05703 -0.0624 1.0000 0.0151
-9.500 -0.5627 0.05755 0.05382 -0.0597 1.0000 0.0151
-9.250 -0.5906 0.05531 0.05152 -0.0546 1.0000 0.0151
-9.000 -0.5937 0.04942 0.04528 -0.0562 0.9884 0.0151
-8.750 -0.5926 0.04326 0.03859 -0.0570 0.9763 0.0158
-8.500 -0.5857 0.03841 0.03308 -0.0568 0.9629 0.0163
-8.250 -0.5739 0.03475 0.02894 -0.0563 0.9482 0.0166
-8.000 -0.5566 0.03171 0.02555 -0.0560 0.9327 0.0168
-7.750 -0.5340 0.02933 0.02288 -0.0564 0.9166 0.0172
-7.500 -0.5069 0.02724 0.02048 -0.0572 0.8999 0.0176
-7.250 -0.4781 0.02542 0.01834 -0.0580 0.8815 0.0181
-7.000 -0.4489 0.02385 0.01648 -0.0587 0.8626 0.0187
-6.750 -0.4211 0.02272 0.01508 -0.0589 0.8439 0.0200
-6.500 -0.3946 0.02158 0.01364 -0.0587 0.8270 0.0209
-6.250 -0.3689 0.02049 0.01232 -0.0583 0.8119 0.0213
-6.000 -0.3439 0.01957 0.01119 -0.0578 0.7982 0.0217
-5.750 -0.3209 0.01842 0.00996 -0.0571 0.7858 0.0224
-5.500 -0.2985 0.01764 0.00910 -0.0564 0.7746 0.0232
-5.250 -0.2762 0.01702 0.00838 -0.0555 0.7645 0.0242
-5.000 -0.2536 0.01651 0.00779 -0.0548 0.7550 0.0255
-4.750 -0.2306 0.01606 0.00723 -0.0541 0.7469 0.0275
-4.500 -0.2082 0.01557 0.00664 -0.0532 0.7394 0.0288
-4.250 -0.1858 0.01508 0.00607 -0.0524 0.7325 0.0308
-4.000 -0.1623 0.01471 0.00563 -0.0518 0.7255 0.0338
-3.750 -0.1385 0.01437 0.00518 -0.0512 0.7196 0.0389
-3.500 -0.1152 0.01395 0.00482 -0.0505 0.7133 0.0559
-3.250 -0.0936 0.01337 0.00446 -0.0498 0.7081 0.1144
-3.000 -0.0746 0.01252 0.00410 -0.0489 0.7033 0.2421
-2.750 -0.0640 0.01103 0.00377 -0.0466 0.6982 0.5059
-2.500 -0.0460 0.01077 0.00417 -0.0440 0.6937 0.6859
-2.250 -0.0218 0.01100 0.00440 -0.0427 0.6895 0.7380
-2.000 0.0011 0.01144 0.00489 -0.0408 0.6846 0.7750
-1.750 0.0246 0.01180 0.00519 -0.0392 0.6804 0.7983
-1.500 0.0517 0.01190 0.00519 -0.0388 0.6770 0.8027
-1.250 0.0788 0.01194 0.00514 -0.0387 0.6733 0.8076
-1.000 0.1060 0.01193 0.00503 -0.0388 0.6691 0.8132
-0.750 0.1331 0.01198 0.00502 -0.0386 0.6652 0.8163
-0.500 0.1604 0.01203 0.00499 -0.0385 0.6620 0.8202
-0.250 0.1877 0.01206 0.00495 -0.0386 0.6587 0.8250
0.000 0.2149 0.01208 0.00494 -0.0386 0.6550 0.8291
0.250 0.2420 0.01212 0.00495 -0.0385 0.6515 0.8320
0.500 0.2695 0.01216 0.00493 -0.0386 0.6483 0.8350
0.750 0.2973 0.01221 0.00492 -0.0388 0.6455 0.8384
1.000 0.3248 0.01223 0.00494 -0.0390 0.6420 0.8423
1.250 0.3518 0.01227 0.00499 -0.0390 0.6385 0.8448
1.500 0.3790 0.01232 0.00504 -0.0389 0.6352 0.8474
1.750 0.4065 0.01239 0.00508 -0.0390 0.6325 0.8505
2.000 0.4340 0.01245 0.00514 -0.0392 0.6294 0.8540
2.500 0.4880 0.01256 0.00533 -0.0393 0.6221 0.8600
2.750 0.5150 0.01263 0.00541 -0.0392 0.6187 0.8630
3.000 0.5423 0.01271 0.00551 -0.0392 0.6154 0.8662
3.250 0.5686 0.01277 0.00564 -0.0392 0.6104 0.8697
3.500 0.5955 0.01281 0.00572 -0.0392 0.6054 0.8730
3.750 0.6220 0.01286 0.00579 -0.0390 0.6006 0.8758
4.000 0.6466 0.01289 0.00592 -0.0385 0.5932 0.8794
4.250 0.6732 0.01290 0.00591 -0.0383 0.5862 0.8830
4.500 0.6981 0.01290 0.00601 -0.0379 0.5767 0.8869
4.750 0.7227 0.01290 0.00609 -0.0373 0.5681 0.8900
5.000 0.7470 0.01291 0.00614 -0.0366 0.5585 0.8938
5.250 0.7711 0.01293 0.00626 -0.0360 0.5477 0.8983
5.500 0.7954 0.01295 0.00635 -0.0354 0.5362 0.9024
5.750 0.8184 0.01297 0.00646 -0.0346 0.5235 0.9062
6.000 0.8413 0.01301 0.00658 -0.0337 0.5082 0.9109
6.250 0.8633 0.01307 0.00668 -0.0327 0.4871 0.9162
6.500 0.8821 0.01317 0.00674 -0.0310 0.4527 0.9215
6.750 0.8959 0.01351 0.00688 -0.0285 0.3980 0.9287
7.000 0.9040 0.01419 0.00725 -0.0253 0.3357 0.9365
7.250 0.9119 0.01502 0.00782 -0.0223 0.2817 0.9464
7.750 0.9393 0.01678 0.00922 -0.0192 0.1939 0.9720
8.000 0.9565 0.01763 0.00996 -0.0186 0.1612 1.0000
8.250 0.9656 0.01844 0.01067 -0.0164 0.1382 1.0000
8.500 0.9768 0.01926 0.01142 -0.0146 0.1190 1.0000
8.750 0.9880 0.02012 0.01224 -0.0129 0.1036 1.0000
9.000 0.9991 0.02102 0.01312 -0.0112 0.0910 1.0000
9.250 1.0099 0.02197 0.01406 -0.0096 0.0807 1.0000
9.500 1.0194 0.02304 0.01513 -0.0081 0.0719 1.0000
9.750 1.0304 0.02407 0.01620 -0.0067 0.0648 1.0000
10.000 1.0390 0.02529 0.01744 -0.0052 0.0598 1.0000
10.250 1.0498 0.02642 0.01865 -0.0041 0.0550 1.0000
10.500 1.0579 0.02778 0.02003 -0.0029 0.0508 1.0000
10.750 1.0677 0.02908 0.02139 -0.0019 0.0468 1.0000
11.000 1.0783 0.03034 0.02270 -0.0011 0.0423 1.0000
11.250 1.0859 0.03188 0.02427 -0.0002 0.0386 1.0000
11.500 1.0974 0.03313 0.02563 0.0004 0.0348 1.0000
11.750 1.1054 0.03472 0.02725 0.0011 0.0320 1.0000
12.000 1.1122 0.03644 0.02905 0.0018 0.0299 1.0000
12.250 1.1221 0.03792 0.03065 0.0024 0.0271 1.0000
12.500 1.1292 0.03968 0.03247 0.0029 0.0248 1.0000
12.750 1.1351 0.04160 0.03446 0.0034 0.0224 1.0000
13.000 1.1420 0.04346 0.03644 0.0038 0.0204 1.0000
13.250 1.1469 0.04554 0.03862 0.0041 0.0187 1.0000
13.500 1.1490 0.04798 0.04113 0.0044 0.0173 1.0000
13.750 1.1526 0.05032 0.04361 0.0047 0.0161 1.0000
14.000 1.1555 0.05279 0.04620 0.0048 0.0150 1.0000
14.250 1.1576 0.05543 0.04893 0.0047 0.0142 1.0000
14.500 1.1561 0.05855 0.05212 0.0044 0.0135 1.0000
14.750 1.1570 0.06148 0.05522 0.0042 0.0130 1.0000
15.000 1.1579 0.06451 0.05841 0.0037 0.0122 1.0000
15.250 1.1575 0.06780 0.06184 0.0031 0.0117 1.0000
15.500 1.1570 0.07119 0.06535 0.0022 0.0112 1.0000
15.750 1.1545 0.07497 0.06926 0.0011 0.0108 1.0000
16.000 1.1509 0.07901 0.07344 -0.0002 0.0106 1.0000
16.250 1.1454 0.08348 0.07804 -0.0019 0.0102 1.0000
16.500 1.1384 0.08832 0.08303 -0.0037 0.0100 1.0000
16.750 1.1322 0.09323 0.08812 -0.0057 0.0098 1.0000
17.000 1.1247 0.09853 0.09361 -0.0080 0.0097 1.0000
17.250 1.1160 0.10419 0.09944 -0.0107 0.0096 1.0000
17.500 1.1064 0.11024 0.10568 -0.0137 0.0095 1.0000
17.750 1.0956 0.11669 0.11231 -0.0171 0.0094 1.0000
18.000 1.0833 0.12377 0.11957 -0.0211 0.0093 1.0000
18.250 1.0700 0.13124 0.12721 -0.0254 0.0093 1.0000
18.500 1.0553 0.13930 0.13545 -0.0302 0.0093 1.0000
18.750 1.0385 0.14826 0.14458 -0.0356 0.0093 1.0000
19.000 1.0195 0.15819 0.15467 -0.0417 0.0094 1.0000
|
Polar data table (+)
Polar graphs
<< Back to NREL's S805A Airfoil (s805a-nr)